1. Filed of the Invention
This invention relates to improved combustor arrangements for gas turbine engines and in particular is concerned with control of air flow to combustor zones.
2. Discussion of Prior Art
The invention relates to improved combustor arrangements for gas turbine engines and in particular is concerned with control of air flow to combustor zones.
Gas turbine engines include an air intake through which air is drawn and thereafter compressed by a compressor to enter a combustor at one or more ports. Fuel is injected into the combustion chamber by means of a fuel injector whence it is atomised, mixed with the compressed air from the various inlet ports and burnt. Exhaust gases are passed out of an exhaust nozzle via a turbine which drives the compressor. In addition to air flow into the combustion chamber through the air inlet ports, air also enters the combustion chamber via the fuel injector itself.
Conventional combustors take a variety of forms. They generally comprise a combustion chamber in which large quantities of fuel are burnt such that heat is released and the exhaust gases are expanded and accelerated to give a stream of uniformly heated gas. Generally the compressor supplies more air than is needed for complete combustion of the fuel and often the air is divided into two or more streams, one stream introduced at the front of the combustion chamber where it is mixed with fuel to initiate and support combustion along with the air in the fuel air mixture from the fuel injector, and one stream is used to dilute the hot combustion product to reduce their temperature to a value compatible with the working range of the turbine.
Gas turbine engines for aircraft are required to operate over a wide range of conditions which involve differing ratios between the mass flows of the combustion and dilution air streams. To ensure a high combustion efficiency, it is usual for the proportion of the total airflow supplied to the burning zone to be determined by the amount of fuel required to be burned to produce the necessary heat input to the turbine at the cruise condition. Often the chamber conditions are stoichiometric in that there is exactly enough fuel for the amount of air; surplus fuel is not completely burnt. However because of variability of the cycles and because air and fuel are never completely mixed there are always some oxides of nitrogen and unburnt fuel residues. An ideal air fuel mixture ratio at cruise usually leads to an over rich mixture in the burning zone at high power conditions (such as take-off) with resultant unburnt hydrocarbon and smoke emission. It is possible to reduce smoke emission at take-off by weakening the burning zone mixture strength but this involves an increase in primary zone air velocity which makes ignition of the engine difficult to achieve, especially at altitude.
The temperature rise of the air in the combustor will depend on the amount of fuel burnt. Since the gas temperature required at the turbine varies according to the operating condition, the combustor must be capable of maintaining sufficient burn over a range of operating conditions. Unwanted emissions rise exponentially with increase in temperature and therefore it is desirable to keep the temperature low. With increasingly stringent legislation against emissions, engine temperature is an increasingly important factor, and operating the combustor at temperatures of less than 2100 K becomes necessary. However at low temperatures, the efficiency of the overall cycle is reduced.
It is a requirement that commercial airliners can decelerate rapidly in the case of potential collision. In order to decelerate a gas turbine from high power to low power, the fuel flow to the engine is reduced. Although the reduction in fuel flow is almost instantaneous, the rate of reduction of engine airflow is relatively slow because of the inertia of rotating parts such as turbines, compressors, shafts etc. This produces a weak mixture of fuel and this increases the risk of flame extinction. It is not always easy to relight the flame especially when the combustor is set to run weakly and at high altitude. Because modern combustors invariably operate in lean burn principles in order to reduce oxide of nitrogen emissions, combustors need to be operated as close to the lean extinction limit at all engine operating conditions. If margins are set wide enough to prevent flame extinction then emissions performance is compromised.
Combustion is initiated and stabilises in the pilot zone, the most upstream section of the combustor. Low power stability requires rich areas within the primary zone of the combustor, enabling combustion to be sustained when the overall air/fuel ratio is much weaker than the flammability limit of kerosene. In traditional combustion systems rich regions can occur in the combustor due to poor mixing and poor atomisation resulting in large droplets of fuel being formed.
Conventional gas turbine engines are thus designed as a compromise rather than being optimised, because of consideration of the above mentioned conflicting requirements at different operating conditions. New xe2x80x9cstagedxe2x80x9d design of combustors overcome the problems to a limited extent. These comprise two combustion zones, a pilot zone and a main zone, each having a separate fuel supply. Essentially this type of combustor is designed such that a fixed flow of about 70% enters the combustor at the main zone and about 30% of the air flows to the pilot zone. In such systems the air/fuel ratio is determined by selecting the amount of fuel in each stage. The air/fuel ratio governs the temperature which determines the amount of emissions. Current gas turbine engine trends are towards increased thrust/weight ratios which require the engine to perform at higher operating compression ratios and wider ranges of combustor air/fuel ratios. Future gas turbine combustion systems will be expected to perform at higher inlet temperatures and richer air/fuel ratios. Because there is little variability in the airflow proportions to the main stage and pilot stage the amount of optimisation achievable for each operating condition is reduced. Even these combustor designs will suffer from either high nitrogen oxide and smoke emissions at full power, or poor stability at low power.
It is therefore desirable to improve control of the amount of fuel, air and air/fuel ratio in each combustor zone to reduce the problems of weak flame extinction, emissions of oxides of nitrogen and unburnt fuel at all operating conditions, whilst maintaining good efficiency and performance.
Conventionally, as shown in GB 785,210, this can be achieved by diverting a main airflow flowing through a main conduit into one of two subsidiary conduits by injecting under pressure into the main airflow a controlling air stream. However, this requires a separate compressor which is disadvantageous in terms of cost and weight. Alternatively, GB 1,184,683 discloses a system whereby a suction action is utilised. However, this is achieved by bleeding compressed air out of the engine resulting in a loss of engine efficiency.
It is an objection of the invention to provide enhanced means by which air flow can be controlled.
According to a first aspect of the present invention, a flow controller for supplying air to a combustor comprises conduit and a control port, the conduit including a main section dividing into at least two secondary sections at a junction characterised in that the control port is positioned in the conduit adjacent to the junction and connected to a reservoir; and wherein, in use, a change in the flow rate of a main airflow flowing through the main section of conduit causes a control airflow to flow either in to or out of the control port whereby the main airflow is selectively diverted into one or other of the secondary sections of conduit.
A change in the flow rate of a main airflow results in a change in the static pressure of the main airflow which produces a pressure differential between the conduit adjacent to the port and the reservoir. The pressure differential causes the control airflow until pressure equalisation, the duration of the flow depending, amongst other things, on the size of the reservoir.
In an alternative embodiment, a flow controller for supplying air to a combustor comprises conduit and a control port, the conduit including a main section dividing into at least two secondary sections at a junction characterised in that the control port is positioned in the conduit adjacent to the junction; and wherein, in use, a control airflow flowing either in to or out of the control port causes a main airflow flowing through the main section of conduit to coanda around a surface of the main section whereby the main airflow is selectively diverted into one or other of the secondary sections of conduit. Ideally, the flow controller comprises at least one arcuate surface common to both the main section and a secondary section.
A skilled person would interpret coanda in relation to the coanda effect, the coanda effect being the tendency of a fluid jet to attach to a downsteam surface roughly parallel to the jet axis. If this surface curves away from the jet the attached flow will follow it deflecting from the original direction (Dictionary of Science and Technology, Larousse 1995).
Preferably, the control port is connected to the conduit further upstream of the junction so as to form a control loop.
In a further embodiment, a flow controller for supplying air to a combustor comprises conduit and a control port, the conduit including a main section dividing into at least two secondary sections at a junction characterised in that the control port is positioned in the conduit adjacent to the junction and connected to the conduit further upstream of the junction so as to form a control loop; and wherein, in use, a control airflow flowing either in to or out of the control port causes a main airflow flowing through the main section of conduit to be selectively diverted into one or other of the secondary sections of conduit.
Preferably, the main section of conduit comprises a convergent-divergent duct; wherein, in use, the control airflow flowing either in to or out of the control port is caused by a pressure differential across the duct.
According to a second aspect of the present invention, a gas turbine combustor comprises a flow controller as described above. Ideally, the flow controller comprises two secondary sections of conduit connected to two different zones within the combustor. In a preferred embodiment, the flow controller comprises one secondary section of conduit connected to a pilot combustion zone within the combustor and another secondary section of conduit connected to a main combustion zone.
In this way the proportion of flow to the main combustor zone and the pilot zone can be selectively altered without mechanical means. This provides robust control of flow with high reliability.